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NASA NTRS2007

Integrated System Test Approaches for the NASA Ares I Crew Launch Vehicle

The Ares I Crew Launch Vehicle (CLV) is being developed by the U.S. National Aeronautics and Space Administration (NASA) to provide crew access to the International Space Station (ISS) and, together with the Ares V Cargo Launch Vehicle (CaLV), serves as one component of a future launch capability for human exploration of the Moon. During the system requirements definition process and early design cycles, NASA defined and began implementing plans for integrated ground and flight testing necessary to achieve the first human launch of Ares I. The individual Ares I flight hardware elements: the first stage five segment booster (FSB), upper stage, and J-2X upper stage engine, will undergo extensive development, qualification, and certification testing prior to flight. Key integrated system tests include the Main Propulsion Test Article (MPTA), acceptance tests of the integrated upper stage and upper stage engine assembly, a full-scale integrated vehicle dynamic test (IVDT), aerodynamic testing to characterize vehicle performance, and integrated testing of the avionics and software components. The Ares I-X development flight test will provide flight data to validate engineering models for aerodynamic performance, stage separation, structural dynamic performance, and control system functionality. The Ares I-Y flight test will validate ascent performance of the first stage, stage separation functionality, and a highaltitude actuation of the launch abort system (LAS) following separation. The Orion-1 flight test will be conducted as a full, un-crewed, operational flight test through the entire ascent flight profile prior to the first crewed launch.

Cockrell, Charles E., Jr.↗

NASA NTRS2008

Information Extraction for System-Software Safety Analysis: Calendar Year 2007 Year-End Report

This annual report describes work to integrate a set of tools to support early model-based analysis of failures and hazards due to system-software interactions. The tools perform and assist analysts in the following tasks: 1) extract model parts from text for architecture and safety/hazard models; 2) combine the parts with library information to develop the models for visualization and analysis; 3) perform graph analysis on the models to identify possible paths from hazard sources to vulnerable entities and functions, in nominal and anomalous system-software configurations; 4) perform discrete-time-based simulation on the models to investigate scenarios where these paths may play a role in failures and mishaps; and 5) identify resulting candidate scenarios for software integration testing. This paper describes new challenges in a NASA abort system case, and enhancements made to develop the integrated tool set.

Malin, Jane T.↗

NASA NTRS2008

A Status Report on the Parachute Development for NASA's Next Manned Spacecraft

NASA has determined that the parachute portion of the Landing System for the Crew Exploration Vehicle (CEV) will be Government Furnished Equipment (GFE). The Earth Landing System has been designated CEV Parachute Assembly System (CPAS). Thus a program team was developed consisting of NASA Johnson Space Center (JSC) and Jacobs Engineering through their Engineering and Science Contract Group (ESCG). Following a rigorous competitive phase, Airborne Systems North America was selected to provide the parachute design, testing and manufacturing role to support this team. The development program has begun with some early flight testing of a Generation 1 parachute system. Future testing will continue to refine the design and complete a qualification phase prior to manned flight of the spacecraft. The program team will also support early spacecraft system testing, including a Pad Abort Flight Test in the Fall of 2008

Sinclair, Robert↗

NASA NTRS2007

CEV Trajectory Design Considerations for Lunar Missions

The Crew Exploration Vehicle (CEV) translational maneuver Delta-V budget must support both the successful completion of a nominal lunar mission and an "anytime" emergency crew return with the potential for much more demanding orbital maneuvers. This translational Delta-V budget accounts for Earth-based LEO rendezvous with the lunar surface access module (LSAM)/Earth departure stage (EDS) stack, orbit maintenance during the lunar surface stay, an on-orbit plane change to align the CEV orbit for an in-plane LSAM ascent, and the Moon-to-Earth trans-Earth injection (TEI) maneuver sequence as well as post-TEI TCMs. Additionally, the CEV will have to execute TEI maneuver sequences while observing Earth atmospheric entry interface objectives for lunar high-latitude to equatorial sortie missions as well as near-polar sortie and long duration missions. The combination of these objectives places a premium on appropriately designed trajectories both to and from the Moon to accurately size the translational V and associated propellant mass in the CEV reference configuration and to demonstrate the feasibility of anytime Earth return for all lunar missions. This report examines the design of the primary CEV translational maneuvers (or maneuver sequences) including associated mission design philosophy, associated assumptions, and methodology for lunar sortie missions with up to a 7-day surface stay and with global lunar landing site access as well as for long duration (outpost) missions with up to a 210-day surface stay at or near the polar regions. The analyses presented in this report supports the Constellation Program and CEV project requirement for nominal and anytime abort (early return) by providing for minimum wedge angles, lunar orbit maintenance maneuvers, phasing orbit inclination changes, and lunar departure maneuvers for a CEV supporting an LSAM launch and subsequent CEV TEI to Earth return, anytime during the lunar surface stay.

Condon, Gerald L.↗

NASA NTRS2008

NASA Propulsion Investments for Exploration and Science

The National Aeronautics and Space Administration (NASA) invests in chemical and electric propulsion systems to achieve future mission objectives for both human exploration and robotic science. Propulsion system requirements for human missions are derived from the exploration architecture being implemented in the Constellation Program. The Constellation Program first develops a system consisting of the Ares I launch vehicle and Orion spacecraft to access the Space Station, then builds on this initial system with the heavy-lift Ares V launch vehicle, Earth departure stage, and lunar module to enable missions to the lunar surface. A variety of chemical engines for all mission phases including primary propulsion, reaction control, abort, lunar ascent, and lunar descent are under development or are in early risk reduction to meet the specific requirements of the Ares I and V launch vehicles, Orion crew and service modules, and Altair lunar module. Exploration propulsion systems draw from Apollo, space shuttle, and commercial heritage and are applied across the Constellation architecture vehicles. Selection of these launch systems and engines is driven by numerous factors including development cost, existing infrastructure, operations cost, and reliability. Incorporation of green systems for sustained operations and extensibility into future systems is an additional consideration for system design. Science missions will directly benefit from the development of Constellation launch systems, and are making advancements in electric and chemical propulsion systems for challenging deep space, rendezvous, and sample return missions. Both Hall effect and ion electric propulsion systems are in development or qualification to address the range of NASA s Heliophysics, Planetary Science, and Astrophysics mission requirements. These address the spectrum of potential requirements from cost-capped missions to enabling challenging high delta-v, long-life missions. Additionally, a high specific impulse chemical engine is in development that will add additional capability to performance-demanding space science missions. In summary, the paper provides a survey of current NASA development and risk reduction propulsion investments for exploration and science.

Smith, Bryan K.↗

NASA NTRS2008

The Manufacturing Process for the NASA Composite Crew Module Demonstration Structure

This paper will describe the approaches and methods selected in fabrication of a carbon composite demonstration structure for the Composite Crew Module (CCM) Program. The program is managed by the NASA Safety and Engineering Center with participants from ten NASA Centers and AFRL. Multiple aerospace contractors are participating in the design development, tooling and fabrication effort as well. The goal of the program is to develop an agency wide design team for composite habitable spacecraft. The specific goals for this development project are: a).To gain hands on experience in design, building and testing a composite crew module. b) To validate key assumptions by resolving composite spacecraft design details through fabrication and testing of hardware. This abstract is based on Preliminary Design data..The final design will continue to evolve through the fall of 2007 with fabrication mostly completed by conference date. From a structures perspective, the.CCM can be viewed as a pressure module with variable pressure time histories and a series of both impact and quasi-static, high intensity point, line, and area distributed loads. The portion of the overall space vehicle being designed and. fabricated by the CCM team is just the pressure module and primary loading points. The heaviest point loads are applied and distributed to the pressure module at.an aluminum Service Module/Alternate Launch Abort System (SM/ALAS) fittings and at Main and Drogue Chute fittings. Significant line loads with metal to metal impact is applied at.the Lids ring. These major external point and line loads as well as pressure impact loads (blast and water landing) are applied to the lobed floor though the reentry shield and crushable materials. The pressure module is divided into upper and lower. shells that mate together with a bonded belly band splice joint to create the completed structural assembly. The benefits of a split CCM far outweigh the risks of a joint. These benefits include lower tooling cost and less manufacturing risk. Assembly of the top and bottom halves of the pressure shell will allow access to the interior of the shell throughout remaining fabrication sequence and can also potentially permit extensive installation of equipment and .crew facilities prior to final assembly of the two shell halves. A Pi pre-form is a woven carbon composite material which is provided in pre-impregnated form and frozen for long term storage. The cross-section shape allows the top of the pi to be bonded to a flat or curved surface with a second flat plate composite section bonded between two upstanding legs of the Pi. One of the regions relying on the merits of the Pi pre-form is the backbone. All connections among plates of the backbone structure, including the upper flanges, and to the lobe base of the pressure shell are currently joined by Pi pre-forms. The intersection of backbone composite plates is formed by application of two Pi pre-forms, top flanges and lobed surfaces are bonded with one Pi pre-form. The process of applying the pre-impregnated pi-preform will be demonstrated to include important steps like surface preparation, forming, application of pressure dams, vacuum bagging for consolidation, and curing techniques. Chopped carbon fiber tooling was selected over other traditional metallic and carbon fiber tooling. The requirement of schedule and cost economy for a moderate reuse cure tool warranted composite tooling options. Composite tooling schedule duration of 18 weeks compared favorably against other metallic tooling including invar tooling. Composite tooling also shows significant cost savings over low CTE metallic options. The composite tooling options were divided into two groups and the final decision was based on the cost, schedule, tolerance, temperature, and reuse requirements.

Pelham, Larry↗

NASA NTRS2008

Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle

As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst-case scenario occurs just before Upper Stage Main Engine Cut-Off when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed Simulink model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and its engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the Simulink model vs. a full geometric ADAMS model are discussed as well.

Konno, Kevin E.↗

NASA NTRS2008

Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle

As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst case scenario occurs just before Upper Stage Main Engine Cut-Off (MECO) when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed Simulink model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and its engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the Simulink model vs. a full geometric ADAMS model are discussed as well.

Konno, Kevin E.↗

NASA NTRS2008

Technical Progress on the Ares I-X Flight Test

Ares I-X will be NASA's first test flight for a new human-rated launch vehicle since 1981, and the team is well on its way toward completing the vehicle's design and hardware fabrication for an April 2009 launch. This uncrewed suborbital development test flight gives NASA its first opportunities to: gather critical data about the flight dynamics of the integrated launch vehicle; understand how to control its roll during flight; better characterize the stage separation environments during future flight; and demonstrate the first stage recovery system. The Ares I-X Flight Test Vehicle (FTV) incorporates a mix of flight and mockup hardware. It is powered by a four-segment solid rocket booster, and will be modified to include a fifth, spacer segment; the upper stage, Orion crew exploration vehicle, and launch abort system are simulator hardware to make the FTV aerodynamically similar to the same size, shape, and weight of Ares I. The Ares IX first stage includes an existing Shuttle solid rocket motor and thrust vector control system controlled by an Ascent Thrust Vector Controller (ATVC) designed and built by Honeywell International. The avionics system will be tested in a dedicated System Integration Laboratory located at Lockheed Martin Space Systems (LMSS) in Denver, Colorado. The Upper Stage Simulator (USS) is made up of cylindrical segments that will be stacked and integrated at Kennedy Space Center (KSC) for launch. Glenn Research Center is already building these segments, along with their internal access structures. The active Roll Control System (RoCS) includes two thruster units harvested from Peacekeeper missiles. Duty cycle testing for RoCS was conducted, and fuel tanking and detanking tests will occur at KSC in early 2008. This important flight will provide valuable experience for the ground operations team in integrating, stacking, and launching Ares I. Data from Ares I-X will ensure the safety and reliability of America's newest launch vehicle.

Davis, S.R.↗

NASA NTRS2008

Summary of Liquid Propulsion System Needs in Support of the Constellation Program

In January 2004, the President of the United States established the Vision for Space Exploration (VSE) to complete the International Space Station, retire the Space Shuttle and develop its replacement, and expand the human presence on the Moon as a stepping stone to human exploration of Mars and worlds beyond. In response, NASA developed the Constellation Program, consisting of the components shown in Figure 1. This paper will summarize the manned spaceflight liquid propulsion system needs in support of the Constellation Program over the next 10 years. It will address all liquid engine needs to support human exploration from low Earth orbit (LEO) to the lunar surface, including an overview of engines currently under contract, those baselined but not yet under contract, and those propulsion needs that have yet to be initiated. There may be additional engine needs for early demonstrators, but those will not be addressed as part of this paper. Also, other portions of the VSE architecture, including the planned Orion abort test boosters and the Lunar Precursor Robotic Program, are not addressed here as they either use solid motors or are focused on unmanned elements of returning humans to the Moon.

Lorier, Terry↗

NASA NTRS2008

Summary of Propulsion System Needs in Support of Project Constellation

In January 2004, the President of the United States established the Vision for Space Exploration (VSE) to return man to the moon and ultimately to extend manned space travel to Mars. This paper will summarize the manned space flight liquid propulsion system needs in support of Project Constellation over the next 10 years. It will include all engine needs to return man to the moon. An overview of engines currently under contract, those baselined but not yet under contract, and those engine needs that hav.e yet to be initiated. Project Constellation includes the components as shown Figure 1. Liquid propulsion systems supporting the manned portion of these elements include the following: the Crew Exploration Vehicle named Orion (crew module reaction control system (CMRCS), service module Orion Main Engine (OME), service module auxiliary RCS, and service module reaction control system (SMRCS)), the Crew Launch Vehicle named Ares 1 (J2X upper stage, first stage roll control system, second stage reaction control system, and the Ares I-X roll control system), the Heavy Lift Launch Vehicle named Ares V (RS68B first stage booster, J-2X upper stage, roll control systems, and the Earth Departure Stage (EDS) (powered by the same Ares V Upper Stage J-2X), and the Lunar Lander named Altair with both descent and ascent stages (lunar orbit insertion and descent main engine, ascent main engine, and attitude control systems for both stages). In addition, there may be additional engine needs for early demonstrators, but those will not be speculated on as part of this paper. Also, other portions of the VSE architecture, including the planned Orion abort demonstrations and the Lunar Precursor Robotic Program, are not addressed here as they either use solid motors or are focused on unmanned precursor missions.

Sumrall, Phil↗

NASA NTRS2008

Model-Based Fault Tolerant Control

The Model Based Fault Tolerant Control (MBFTC) task was conducted under the NASA Aviation Safety and Security Program. The goal of MBFTC is to develop and demonstrate real-time strategies to diagnose and accommodate anomalous aircraft engine events such as sensor faults, actuator faults, or turbine gas-path component damage that can lead to in-flight shutdowns, aborted take offs, asymmetric thrust/loss of thrust control, or engine surge/stall events. A suite of model-based fault detection algorithms were developed and evaluated. Based on the performance and maturity of the developed algorithms two approaches were selected for further analysis: (i) multiple-hypothesis testing, and (ii) neural networks; both used residuals from an Extended Kalman Filter to detect the occurrence of the selected faults. A simple fusion algorithm was implemented to combine the results from each algorithm to obtain an overall estimate of the identified fault type and magnitude. The identification of the fault type and magnitude enabled the use of an online fault accommodation strategy to correct for the adverse impact of these faults on engine operability thereby enabling continued engine operation in the presence of these faults. The performance of the fault detection and accommodation algorithm was extensively tested in a simulation environment.

Kumar, Aditya↗

NASA NTRS2008

Simulation Assisted Risk Assessment Applied to Launch Vehicle Conceptual Design

A simulation-based risk assessment approach is presented and is applied to the analysis of abort during the ascent phase of a space exploration mission. The approach utilizes groupings of launch vehicle failures, referred to as failure bins, which are mapped to corresponding failure environments. Physical models are used to characterize the failure environments in terms of the risk due to blast overpressure, resulting debris field, and the thermal radiation due to a fireball. The resulting risk to the crew is dynamically modeled by combining the likelihood of each failure, the severity of the failure environments as a function of initiator and time of the failure, the robustness of the crew module, and the warning time available due to early detection. The approach is shown to support the launch vehicle design process by characterizing the risk drivers and identifying regions where failure detection would significantly reduce the risk to the crew.

Mathias, Donovan L.↗

NASA NTRS2008

Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle

As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst case scenario occurs just before Upper Stage Main Engine Cut-Off (MECO) when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed 2-D (Simulink) model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and its engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the 2-D model vs. a full 3-D (ADAMS) model are discussed as well.

Konno, Kevin E.↗

NASA NTRS2009

Status of Constellation Pressure Garment Development

The Constellation Program requires the development of a space suit system to meet new requirements for launch, entry, and abort crew survival functions, microgravity intravehicular and extravehicular activities, and lunar surface exploration. This paper summarizes recent work and the current status of the NASA Constellation Space Suit Element Pressure Garment and Crew Survival Subsystem (PG/CS). The emphasis of the work by the PGS/CS team has been in the areas of feasibility studies toward PGS/CS architecture definition, risk mitigation, and requirements development. Included are results from component level engineering studies, testing in the Orion Vehicle and Orion seat mockups, microgravity testing on the Reduced Gravity Aircraft, occupant protection sled testing, analyses and studies, and their implications on Constellation PG/CS subsystem.

Ross, Amy↗

NASA NTRS2008

Orbital Express Advanced Video Guidance Sensor

In May 2007 the first US fully autonomous rendezvous and capture was successfully performed by DARPA's Orbital Express (OE) mission. Since then, the Boeing ASTRO spacecraft and the Ball Aerospace NEXTSat have performed multiple rendezvous and docking maneuvers to demonstrate the technologies needed for satellite servicing. MSFC's Advanced Video Guidance Sensor (AVGS) is a primary near-field proximity operations sensor integrated into ASTRO's Autonomous Rendezvous and Capture Sensor System (ARCSS), which provides relative state knowledge to the ASTRO GN&C system. This paper provides an overview of the AVGS sensor flying on Orbital Express, and a summary of the ground testing and on-orbit performance of the AVGS for OE. The AVGS is a laser-based system that is capable of providing range and bearing at midrange distances and full six degree-of-freedom (6DOF) knowledge at near fields. The sensor fires lasers at two different frequencies to illuminate the Long Range Targets (LRTs) and the Short Range Targets (SRTs) on NEXTSat. Subtraction of one image from the other image removes extraneous light sources and reflections from anything other than the corner cubes on the LRTs and SRTs. This feature has played a significant role for Orbital Express in poor lighting conditions. The very bright spots that remain in the subtracted image are processed by the target recognition algorithms and the inverse-perspective algorithms, to provide 3DOF or 6DOF relative state information. Although Orbital Express has configured the ASTRO ARCSS system to only use AVGS at ranges of 120 m or less, some OE scenarios have provided opportunities for AVGS to acquire and track NEXTSat at greater distances. Orbital Express scenarios to date that have utilized AVGS include a berthing operation performed by the ASTRO robotic arm, sensor checkout maneuvers performed by the ASTRO robotic arm, 10-m unmated operations, 30-m unmated operations, and Scenario 3-1 anomaly recovery. The AVGS performed very well during the pre-unmated operations, effectively tracking beyond its 10-degree Pitch and Yaw limit-specifications, and did not require I-LOAD adjustments before unmated operations. AVGS provided excellent performance in the 10-m unmated operations, effectively tracking and maintaining lock for the duration of this scenario, and showing good agreement between the short and long range targets. During the 30-m unmated operations, the AVGS continuously tracked the SRT to 31.6 m, exceeding expectations, and continuously tracked the LRT from 8.8 m out to 31.6 m, with good agreement between these two target solutions. After this scenario was aborted at a 10-m separation during remate operations, the AVGS tracked the LRT out 54.3 m, until the relative attitude between the vehicles was too large. The vehicles remained apart for eight days, at ranges from 1 km to 6 km. During the approach to remate in this recovery operation, the AVGS began tracking the LRT at 150 m, well beyond the OE planned limits for AVGS ranges, and functioned as the primary sensor for the autonomous rendezvous and docking.

Howard, Ricky↗

NASA NTRS2008

Improved Verification for Aerospace Systems

Aerospace systems are subject to many stringent performance requirements to be verified with low risk. This report investigates verification planning using conditional approaches vice the standard classical statistical methods, and usage of historical surrogate data for requirement validation and in verification planning. The example used in this report to illustrate the results of these investigations is a proposed mission assurance requirement with the concomitant maximum acceptable verification risk for the NASA Constellation Program Orion Launch Abort System (LAS). This report demonstrates the following improvements: 1) verification planning using conditional approaches vice classical statistical methods results in plans that are more achievable and feasible; 2) historical surrogate data can be used to bound validation of performance requirements; and, 3) incorporation of historical surrogate data in verification planning using conditional approaches produces even less costly and more reasonable verification plans. The procedures presented in this report may produce similar improvements and cost savings in verification for any stringent performance requirement for an aerospace system.

Powell, Mark A.↗

NASA NTRS2008

Method and associated apparatus for capturing, servicing and de-orbiting earth satellites using robotics

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Cepollina, Frank J.↗